This invention relates to air inlets for aircraft turbine engines and is directed more particularly to an air inlet for the turbine engine of a subsonic aircraft.
Boundary-layer bleed systems are known and have been used to improve the performance and control of supersonic inlets where the problems of shock-boundary layer interactions and supersonic diffusion are encountered. The diffusion of supersonic flow has not been usually associated with subsonic inlets. However, supersonic or high Mach number diffusion can arise during subsonic flight if the inlet is subjected to a sufficiently high combination of forward velocity and angle of attack. This causes boundary layer flow separation in the inlet at the windward side with a resultant reduction of pressure recovery and increased distortion of the air flowing from the inlet into the turbine engine. Such situations have been encountered in inlets for VTOL and STOL aircraft in the departure and approach portions of flight. The inlets for highly maneuverable military aircraft may also encounter similar problems.